1. Technical Field
This invention relates generally to aircraft gas turbine engines and particularly to the cooling of the platform sections of turbine airfoils employed in such engines.
2. Background Art
The operation of gas turbine engines is well known. Such engines include a serial arrangement of a fan, a compressor, a combustor and a turbine. Air admitted into the inlet of the engine is compressed by the engine's compressor. The compressed air is then mixed with fuel in the engine's combustor and burned. The high-energy products of combustion of the burned air-fuel mixture then enter the turbine which extracts energy from the mixture in order to drive the compressor and fan. That energy extracted by the turbine above and beyond that which is necessary to drive the compressor and fan exits the engine at the core engine exhaust nozzle thereof producing thrust which powers an associated aircraft or operates a free turbine which drives an electrical generator, pump or the like.
As gas turbine engines evolve, they have been required to produce greater and greater quantities of thrust, often resulting in higher engine operating temperatures and higher stresses in various engine components, particularly turbine blades. The combustor temperatures of modern high performance gas turbine engines often exceed the melting temperature of the materials from which the turbine blades are manufactured. Therefore, such blades must be cooled, usually by air bled off the engine's compressor. The blades are typically provided with internal cooling passages extending therethrough. Cooling air passing through the cooling passages keeps the blade cool enough to prevent the melting thereof by the high temperature combustor gases. In many respects, cooling turbine blades in this manner has been effective at minimizing oxidation, creep, and thermo-mechanical fatigue, particularly, in the airfoil sections of such blades, due in large measure to the airfoil's ability to accommodate complex networks of intricate patterns of cooling passages.
While turbine blade configurations may lend themselves to cooling airfoils in such manner, such is not necessarily the case with the blade platforms. Such platforms tend to take the shape of longitudinally and circumferentially extensive thin plates which are not conducive to the provision therein of internal cooling passages. In fact, in many instances, modern blade platforms employ no internal cooling passages at all, except perhaps for a series of cooling holes extending through the platform between the radially inner and outer major surfaces thereof. Thus, to cool such platforms, particularly in areas thereof remote of the juncture of the platform with the airfoil, it has been the practice to bathe the underside of the platform with compressor bleed cooling air and channeling the cooling air through the cooling holes to the radially outer surface of the platform where the air mixes with the engine's working fluid. In typical gas turbine engine configurations, the air which bathes the underside of the platform is, for the most part, stagnant, the flow of cooling air through the holes resulting from the difference in pressure between the cooling air and the combustion. While bathing the underside of the blade platform in stagnant compressor blade cooling air does provide some cooling, it has been observed that often, such cooling hindered by the presence of platform-to-platform seals which bear against the underside of the platform and may be insufficient to prevent oxidation, creep, and thermo-mechanical fatigue of the platform. The portion of the blade platform near the cooling holes has been particularly susceptible to thermo-mechanical fatigue which manifests itself in cracking in the high stress around the cooling holes.